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Will a Secondary Surveillance Radar be affected by a 300m high wind turbine?
60 wind turbines are located at a distance of about 30 km away from a Secondary Surveillance Radar installation; what will be the impact of this?
My less than stellar knowledge of radar systems tells me that there shouldn't be an issue, since its a secondary radar and not a primary radar (the blades are big enough to be mistaken for an aircraft) which relies on reflection.
But the 30 km distances makes me doubt if this will actually be the case, might a 300m high obstacle in a radar's direct line of sight might cause problems?

It may or may not affect the SSR and it may not matter. As usual: it depends. A few question that pop up in my radar mind: How high is the radar antenna (above reference level,  e.g. MSL or WGS84)? How high is the base of the turbine, how high is the hub and what is the diameter of the rotor. How many blades does the turbine have? Is there line-of-sight? What is the operational range of the SSR? Is there regular air traffic in the shadowed area? Is there a far field monitor in the general direction of the turbines?  Without an answer to these questions, the answer to your question is: maybe.

Wind turbine length: 225m, wind turbine diameter: 150m, 48 blades, direct line of light. operational range 250NM. yes theres regular traffic over the area.

48 blades? Usually there are 3. Sometimes 2. I've seen very small ones with 5. But 48? That's almost a solid disk... Are you sure?

No they won't, since secondary surveillance radar (SSR), as you point out, does not rely on reflecting electronic pulses, but rather on digital communication with an aircraft transponder.
Transponder signals do still travel in a straight line (line of sight), so obviously putting a very large obstacle (such as a mountain) close to an SSR will limit the radar coverage in the area behind the obstacle.

I invite the downvoter to please comment with any factual errors in my answer so that I can correct them, or alternatively post a better answer themselves.

"Transponder signals do still travel in a straight line" You are minimizing the multipath problem which is essential in radio-communications (refraction), and the Fresnel ellipsoid required clearance (due to diffraction). Doppler shift due to turbine rotor relative speed close to aircraft relative speed may also be something to deal with (SSR receiver must compensate Doppler shift due to aircraft movement). (I'm not the downvoter)

@mins Sure, there's a lot more to it (and frankly I'm not qualified to get into all those details). But I think we can agree that - at least from a user (ATC) perspective - some wind turbines 30 km from a radar station is not going to have any noticeable impact on the coverage.

I'll agree it's not a significant impact on SSR today. It's with certainty a problem for PSR, even at that distance (clutter), and possibly one for SSR (sensitivity loss) in particular conditions. It could become a problem for SSR in the future, with increasing low altitude UAV flights.

See also this document from Eurocontrol, where the rational for a recommended 16 km distance between wind turbines and SSR is explained (annex D, page 59).

@mins Maybe you would consider posting a more detailed answer, since you know a lot more about the technical aspects than I do? :)

Yes, for planes whose ground vector intercepts any part of the obstacle. By ground vector I mean line-of-sight from secondary antenna to the air plane. Atmospherics may shift the radiation path, and missing the target on one rotation of the antenna likely wouldn’t degrade the confidence factor, except in unusual circumstances, such as a hovering helicopter or a plane blocked by the wind turbine for 3 or more rotations of the antenna.


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In a Vmc demo, why is it important to lose 1 knot per second?
I got asked two questions on my checkride:

Why is it important to lose 1 knot per second (approximately) in a Vmc Demo?
How do we know we're losing 1 knot per second?

My answer to question 1 would be be: because you are flying with one engine simulated inoperative and creating asymmetrical yaw, if you lose airspeed too fast you might blow past the stall indicator and actually stall the aircraft with a large yawing effect throwing you into a spin. Does that make sense?
For question 2, other than counting 5-10 seconds while losing 5-10 knots of airspeed is there any other way?

Why do we want to lose 1 knot per second in a VMC demo?

To more accurately determine where VMC is:
As Carlo Felicione's answer says, VMC changes day by day, and you want to determine it precisely.  You can't do that if you're slowing too rapidly.

To permit smooth control of the aircraft:
During a VMC demo, rudder required is determined by aircraft speed. The slower you go, the more rudder you need.  The more abruptly your speed changes, the more abruptly your rudder needs to adjust to compensate.  Since the point of a VMC demo is to determine when your rudder fails to control your heading, abrupt speed changes should be avoided.

To allow consistent control techniques:
Have you noticed that it's really hard to hold a particular vertical speed using the vertical speed indicator?  Usually, you select a pitch, hold it, and then reference the VSI to determine whether a pitch adjustment is necessary. The same is true of airspeed loss.  During a VMC demo, usually, the pilot will configure the airplane for engine-out flight and then pitch to an attitude that will create a 1 knot per second loss of airspeed. As pitch settings are fairly unchanging from flight to flight, this allows you to do the maneuver with more consistency.  You can put the rate of airspeed loss more or less in the back of your mind and concentrate on your rudder control, which is the point.

How do we know we're losing 1 knot per second?

Counting: As you say, looking at the airspeed indicator and counting seconds is a simple and reliable way to do it.

Known pitch settings: As mentioned above, pitch to a known setting and you'll get pretty close to 1 knot per second without the fuss of counting.  As you gain more experience this will become the way you'll do it.

Because in a Vmc demo, you are attempting to determine exactly where Vmc is for your airplane in a particular configuration and the way that’s determined is to maintain straight and level (or coordinated, if your light twin cannot maintain altitude on a single engine) flight with the critical engine out and determine the exact airspeed in which you can no longer maintain directional control of the vehicle.  Remember Vmc (redline) is an aircraft design, not a operating, requirement and varies depending on a number of factors, so it may or may not occur at the indicated red radial line on your airspeed indicator, depending on circumstances.  Therefore, we creep toward it, slowly approaching actual Vmc by slowing 1 knot or so of airspeed a second until we can no longer maintain directional control with the rudder.

source: Stack Exchange

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/

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Why is the angle of attack not proportional to the stagnation point?
While I was studying my PPL today, I've encountered a question right below and I wasn't really satisfied with the answer.
Which statement about the airflow around an aerofoil is correct if the angle of attack increases?

The stagnation point moves up
The stagnation point moves down (correct answer)
The center of pressure moves down
The center of pressure moves up

Up here, α is the degree between the chord line and the oncoming airflow.

1) Stagnation point
2) Transition point
3) Transition point
4) Separation point
So, knowing all of these facts I just can't understand the question above. Why is the stagnation point moves down instead of going up while increasing the angle of attack?

If camber goes up? Camber is the curvature of the wing. It goes up when you extend slats or flaps. Angle of attack is the α in the first picture, and goes up by simply rotating the wing upwards.

That's what I said in a simple way or whatever I'll just edit that out thanks for the notification.

From this site:

Notice that at low AoA the stagnation point is smack on the nose of the wing profile, while at higher AoA it is situated on the lower wing surface.
Hence: stagnation point location lowers with increasing AoA.

Leading-edge stagnation point moving down with increasing angle of attack is true for inviscid flows, however the same does not always hold for viscous (or more realistic) flows, especially around stall. For a given airfoil, if angle of attack increases, the leading-edge stagnation point could move up or down from the leading-edge depending on the flow state on the airfoil upper and lower surface.
As an empirical example, Gregory and O’Reilly mapped the pressure distribution around the leading-edge of a NACA 0012 airfoil for a wide range of angles through stall and captured the motion of the leading edge stagnation point 1 with maximum pressure coefficient. Notice how the leading edge stagnation point moves back towards the leading edge after 17 degrees (after stall).

N. Gregory and C. O’Reilly, “Low-Speed Aerodynamic Characteristics of NACA 0012 Aerofoil Section, including the Effects of Upper-Surface Roughness Simulating Hoar Frost,” R&M 3726, Aeronautical Research Council, 1970.

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/


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How much do a DC-9's wings weigh?
How much do the wings weigh?  How much does the landing gear weigh? How much fuel does it take to fill the tanks? We, at Confederation College are trying to donate a DC-9 for training purposes and require information to facilitate this.

Have you tried searching for any of this information?  What did you find?

@dalearn I reckon it is a valid question. Data of weights of parts of structures is not easy to find, especially not of a type manufactured in the pre-search-engine days.

From this answer, for the DC9-10 variant:

the wing group weighs 9470 lbs
the landing gear weighs 3660 lbs

Wikipedia mentions the fuel weight as 24,743 lbs.
For other variants, scale the wing and landing gear weights up with the ratio of Empty Weights found on Wikipedia.


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Why do some airplanes have a vertical rudder hinge while some others are canted?

I was thinking that canted rudder hinges are used for modern aircraft while the vertical ones are used for old slow airplanes. But it seems that I was wrong since the SST Tupelov Tu-144 and the Concorde also have that vertical (or almost vertical) rotational rudder axis. Then my question is, what is the advantage and this advantage of such a rudder? Which one is better?

The vertical tail of the A380 in the last photo is swept back. Surface sweep is used for high subsonic wings and tail surfaces, to postpone the effects pf compressibility when approaching the speed of sound. And when the vertical tail is swept back, the rudder hinge will be at an angle as well: the rudder chord will be a relatively constant percentage of the vertical tail chord, for aerodynamic and structural reasons.
The two photos of supersonic aeroplanes show vertical tails that are not swept back - they cruise faster than the speed of sound and do not have to postpone any compressibility effect, they fly right through it. No reason to sweep anything back here.
And the photo of the little Cessna with swept back vertical tail: it only looks as if it is trying to postpone any compressibility effects. Of course, it does so by flying way slower than the speed of sound, and the vertical tail is only swept back because it looks super fast.

Sweep helps to widen the beta range in which the vertical works well. It is not only for looks.

Having flown a number of straight tail and swept tail single engine Cessnas, I can tell you the effect is not detectable on a GA airplane.  Cessna went to swept tails in the 60s purely for styling reasons.

@qqjkztd Beta here means sideslip angle. Using beta for both airflow direction and propeller pitch is rather silly, given how many greek letters are available ...


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Which source(s) provides the most authorative and correct flight dynamics theory?
Just like many curious people, the question: " How airplanes actually work?" occupied my mind for a long time.
Different sources gave different answers where most claimed the Bernoulli Principle to be the main reason while many disagreeing with that and so on.
Then I came across this: https://physics.stackexchange.com/q/51503/266050
This has many answers and all in great detail. This gave me a lot of knoweldge and pretty much outlined the many factors that are responsible.
Unfortunately, there is no unanimous consensus even on those answers. The accepted answer (like many others) is detailed and beautiful but there are comments disagreeing with that and people holding different views on which factors actually contribute primarily for the working of airplanes.
Now, usually the answers and discussion given on stackexchange would be more than sufficient but in this case what bugs me is that airplanes are actual things used daily and there are people who design and create them. These people (aircraft engineers, scientists etc) surely must be really certain about how these things work (or so I hope ?) and not be confused with things like whether the upper part of the wing pushes air down or actually pulls it up (not that they would have different outcomes or effects but one ought to know!) or whether the angle of attack really does play a role or not. The reason for my assumption of flight engineers having surity of knowledge to the core is that they need to know with high accuracy and precision the conditions and limits under which airplanes will work and when a certain principle would fail.
Why is it then that not a clear concise unanimous theory is available (or is it?) open-source etc that covers every factor and the extent to which they matter and which is accepted (or rather reffered to) unanimously without confusion?
If, on the other hand, there is a source (or sources) available after reading which I can confidently say that I know with very high certainty and with all technicalities how airplanes work, I hope to know where and how I can find it.

I'm voting to close this as opinion based. As you say there isn't a unanimous view, and asking why there isn't one isn't going to get fact based answers. Also, asking for a reading list of 'best' sources is also asking for opinion.

I think you're confusing the existence of multiple descriptions for disagreement (and disagreement among random people on the internet for lack of academic consensus). Each description has different applicability and level of detail, and all descriptions are ultimately a statistical simplification of the underlying quantum soup. Just find a description that is useful for you.

@Sanchises Finding a description useful for me is 'only useful" if its correct. As I have emphasized, the problem is not that people are giving different descriptions of the same thing. They are telling different things altogether.

@GdD I do not see any other site best fit for this question. This is the most basic and important one too. This is 'AviationStackExchange'  and I am simply asking 'where can I find how airplanes work?' I don't know an answer to this since there are so many disagreements. How can this be an opinionated thing? The working of an aircraft is factual.  Isn't it?

@Sanchises Regarding your last couple of lines related to descriptions being 'statistical simplification of quantum soup'. I am not looking for the microscopic detailed analysis. I am simply asking whether there is a consensus theory of working an aircraft. For instance,  If I ask you "what makes an airplane fly?" And you tell me it's Bernoulli and some one else comes along and says its actually primarily due to angle of attack and other says something else. Then I am confused. This doesn't even touch quantum statistics or anything like that. Now, 'disagreement among random people on internet'

That I understand and hence I was asking for credible or authoritative sources. Nonetheless I would say the answers posted on Physicsstackexchange (link that I posted) give references etc and are answered by credible people. This is precisely the reason why it's so bugging.

"Finding a description useful for me is 'only useful" if its correct." well then unfortunately you are going to be disappointed because ALL theories are incorrect.  i.e. all theories necessarily make simplifying assumptions and approximations, and thus are not correct. But many theories are useful. This is not just aerodynamics. E.g. Newton's law of gravity is "incorrect" but it is very useful in many situations.

The title states Flight Dynamics, the link to the Physics site is one about aerodynamics. That seems more relevant to the question and I've added the tag.

There is unanimous agreement (at least among people who actually understand physics) that

the Navier-Stokes equations are the best we've got and
you can't get simpler than that!

By the second point I don't mean it's simple. Not only they are not simple, but they exhibit such complex behavior that it wasn't fully explored even from the pure mathematical standpoint.
What I mean is that all the properties described by the equations (conservation of mass, momentum and energy, viscosity and pressure) are important in creation of lift and you can't skip any of them and still arrive at something at least somewhat approximating reality.
There are some theories with which would can get fairly accurate estimates of lift like the lifting-line theory and the thin airfoil theory with much less calculation, but they all have serious limitations and depend on experimentally-determined factors. They don't really explain lift itself, only some relations around it like what is effect of increasing span or chord.

"you can't skip any of them and still arrive at something at least somewhat approximating reality" Okay this somewhat clears some things up. Though tell me, is there a factor or phenomenon without which lift will be not possible or sufficient?

And by a factor or phenomenon, I mean something that is engineered into the architect of the aircraft which is so crucial without which the aircraft wouldn't lift even while other factors were present. For example, from my limited knowledge, I understand that there is a critical angle of attack which produces max lift and that the wings have a larger surface area for Bernoulli upthrust. Now, of course, both are utilized but lets suppose one of these factors were removed could the aircraft could still lift?

@Lost You can produce lift with other devices than airfoils, for example spinning cylinders (Magnus effect), but you can only do so in a fluid that has both mass and viscosity. For airfoils without viscosity the rear stagnation point wouldn't stick to the trailing edge and wouldn't be turning the air downward (the Kutta condition wouldn't satisfied). And without mass there would of course be nothing to give the momentum to.

@Lost For an airfoil, the most important feature is its sharp trailing edge. A flat plate at an angle to the stream will produce lift, and a slightly bent plate is as good at it as more complex shapes at its optimal angle of attack. But without sharp trailing edge, the air from below will bend around it, cancelling the downward slope of the wing and that will prevent generating lift. On the other hand the leading edge needs to be blunt to better split the air at different angles to keep drag low.

A complete theory is taught in universities and takes at least 5 years to master. You would know quite a lot of the science of aerodynamics after obtaining your doctorate, which takes even longer.
Or you could obtain a degree in aeronautical engineering. That is where you learn about the multiple fields of technology that keeps an aircraft in the air and enables passengers to trust them with their lives when flying to Bali. But now there's an issue: there are way more subjects to be taught in 5 years than only fluid dynamics, like thin walled structures, systems technology, safety engineering, human-machine-interaction etc, plus the mathematics and physics that lies at the foundation of all of this. So now we need to start making judgements on where to stop.
What to teach and what not is a function of the end goal and the audience.

The design engineers of a passenger aeroplane needs to be able to integrate all relevant knowledge to an extent that enables people flying to Bali to trust the machine with their lives.
Focusing on pilot training: quick decision making and rapid and consistent control is the key element. The airplane is already there, it is  tested in the wind tunnel, constructed in the factory, flight tested and calibrated by the test pilots etc. So it is important to provide quick understanding of the principles of flight, only relevant to keeping the plane in the air.

So why do wings provide lift? For a pilot, all knowledge of cruise, max. altitude, static and dynamic stall, anything to do with adverse aircraft state and navigation is the main objective here - we leave out what is not of direct importance, and move on to the science of wind shear at approach. The comments on the physics site comment on what is left out of a relatively simple explanation that is sufficient for 80% of cases.
There is full agreement on what is correct aerodynamics theory. Now with computers, we can finally use the Navier Stokes equations to compute aerodynamics acting on wings. But a very important decision remains: who is the audience, and what would best be left out.

From the answers, it seems that in order to exactly understand what actually plays the most significant role one would have to wait till Navier-Stokes gets solved. And all the engineering and Physics that goes into building an aircraft is a continuous process of experimental refinement. Making a statement like "Bernoulli's or X is what makes the airplane fly" is highly incomplete because we don't know how much role, if any, it plays. We just know after many trials and errors that "shaping the wing and nose etc in a particular way" gives max lift. Am I thinking about this correctly?

There are many “computer fluid dynamics” software packages that solve Navier-Stokes equations, and they are being used when designing aircraft, because it's faster and cheaper than making lots of different models and testing them all in awind tunnel. The trouble is that it's a long way from a set of differential equations to a some kind of insight. But you can get a bit of that on the quantitative level. The linked Q/A already does explain that.

@Lost Note that there is no or here. Bernoulli's equation is part of the Navier-Stokes set. After all it's just the law of conservation of energy for fluids (usually given for incompressible flow, but for higher speeds the compressible form is needed, which additionally requires pulling in a bunch of themrodynamic equations to make up for the additional free variables). It is, however, just one equation with multiple free variables, so you need additional equations to get a solution. And no, it does not take the air the same time to get around the wing—the air above makes it much faster.

Your question may not have a certain and definitive answer.  Nevertheless, there are two particular areas (and two particular books) that can be discussed in guiding your attempt to find a definitive answer.  Indeed, what is an authoritative source for understanding flight-dynamics theory?
The two areas are -

The principles of flight
The elements of aircraft design

These two areas are an inseparable and expansive body of knowledge.  For the engineering practitioner, the common core areas linking these two areas are many, requiring an understanding of a) mathematics at least through differential equations (matrix theory and linear algebra are especially helpful), b) the principles of physics and atmospheric science, and c) engineering and applied science.  Within this last area should be knowledge particularly in disciplines of d) fluid mechanics, e) strength of materials and structures, and f) applied mechanical engineering, particularly in g) materials science, h) power and propulsion systems, and i) aircraft systems design.  All of these areas are inextricably linked through experimentation and assessment.  Properly defining a problem requires a complete understanding of the problem. Things we think of as authoritative and complete are rarely, if ever, totally comprehensive and inclusive or exhaustive. Flight dynamics theory is as much an attempt to understand why things work, as to understand why things don't work.
For a start, the following is suggested -
Gudmundsson, Snorri, 2014, General Aviation Aircraft Design: Applied Methods and Procedures.  Butterworth-Heinemann (Elsevier), Oxford. 1034 pp, several appendices.
Upon examination of this book and others, you may be given pause; "but wait... this is just what I did not want to do!"  So lets look at another book that is a superbly easy read, absolutely indispensable, and requires only algebra and a little calculus to understand.  That book is -
Dommasch, Daniel O., Sherby, Sydney S., and Connolly, Thomas E., 1967, Airplane Aerodynamics, 4th ed. Pitman Publishing Corp., New York.  621 pp.
This book has a clear and authoritative discussion of the principles of flight, and an overview of factors involved in aircraft aerodynamics and flight performance.  When you read through this book, you will get a pretty good idea regarding the scope of flight dynamics without frying your brain.  This text is particularly helpful regarding propeller driven aircraft.  At the heart of the book, which is easily understood, are two central sections, 11:7 and 11:8, regarding the V-g diagram.  All of the factors regarding why airplanes break are discussed therein.  Gudmundsson's book has an in-depth engineering discussion of the same topic starting in section 16.4; only algebra is needed to work through his examples.
But you may say, I just want to know why airplanes fly.  Most books you will read will discuss the part regarding how airplanes fly, leaving the why as seeming conceptually difficult or lacking.  When you find references to theory, you should pay close attention (even if your brain gets fried) because that is where the conceptual developments and basic understanding are found; sometimes difficult, as they may be.
A suggestion would be to read, read, read.  Take lots of notes, write your own personal text, capture your own ideas, maybe experiment with the math.  And look at the cited references in what you read; these may be real gems.  See what you think. This can be fun...


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How to determine the thickness of an airfoil?
In calculation or designing, how to determine the airfoil thickness? What is the guideline or consideration in designing airplane's wing? We all know that the bulge (the ticker in the airfoil) is required to apply Bernoulli principle, to create low pressure above the wing. But how to bind the round nose shape requirement to the design requirement. In this case, to make the airplane to be efficient (especially related to the power consumption).
This airfoil is required for an airplane as I explained here. Maximum weight will be 400kgs, and the maximum speed will be 200km/h.

We all know that the bulge (the ticker in the airfoil) is required to apply Bernoulli principle, to create low pressure above the wing. You may want to read up on that assumption. See related questions  on airfoil thickness, equal transit time fallacy and how wings generate lift

Sorry, I may have misunderstood what you meant with "bulge". I edited the question and replaced "bulge" with thickness, is that what you mean or do you mean Nose Shape?

Factors for determining the wing thickness during the aeroplane design procedure:

Determine the required structure height, which follows from the aspect ratio determined by the aerodynamicists: the cantilever ratio, which is the structural wing semi-span divided by maximum root thickness. Statistical analysis has found that transport aircraft usually have a ratio between 18 and 22, lower means a lighter wing but more drag.
Stall behaviour: for thin wings the leading edge stall dominates, for thick wings trailing edge stall which is more gradual. This imposes a minimum thickness.
For aircraft cruising at M > 0.5: the wing thickness must be such that dive Mach number remain subcritical. Imposes a maximum thickness.
For good max $C_L$ a value from the sweet spot between 12 and 15 should be selected, see below = fig 7-9 from Torenbeek.

Torenbeek then concludes:

It is concluded that the root section thickness of transport aircraft should be chosen such that a good cantilever ratio is obtained. For a given taper ratio this implies that the thickness ratio at the root will increase proportionally to the aspect ratio. A thickness ratio of between 15 and 20 percent is in the interest of good performance when using relatively simple trailing-edge high-lift devices and provides adequate room for retracting the undercarriaqe. Thickness ratios above 20 percent may show diminishing returns due to the increasing profile drag and the relatively low maximum lift and this, in turn, limits the aspect ratio to a maximum of approximately 13 for cantilever wings.
Tip sections (without flaps) should be between 10 and 15 percent in order to attain a high maximum lift. This reduction relative to the root is also in favor of low structural weight.

The conclusion of Item #2 and #3 aren't swapped? I guess the #2 imposes thicker airfoil and #3 imposes thinner airfoils. Maybe I'm wrong, of course.

@AlexByrth you’re exactly right. Imposing a thicker airfoil means there is a minimum value, below which there would be problems.

thanks for your explanation. I'm a non-native English speaker, so I've translated "impose" as "force to a", "obligate to a". But seems it means "limits the value of".


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What happened to this plane during landing? Failed first time?
Hi I was checking flights at Heathrow on flightradar24.com the other day and found this plane a bit strange. It looked like the first landing wasn't successful. The plane may have touched down but pulled up again, and then came back for a 2nd landing. Does anybody have any factual information on what happened? Thanks.

AT800 / RAM800F, 2nd April 2022, landing

There are many many reasons why an aircraft might go around or perform a missed approach (in fact it is often said that pilots plan to go around, and a landing is welcome but unexpected). Try googling those terms for more background. Without listening to a recording of the radio conversation at the time, it would be very difficult to say what the exact reason is for this exact go-around.

Same day, a british airways flight did the same https://www.youtube.com/watch?v=jMGsm89GYSs&t=7645s

There are many, many reasons why a pilot might go around.

The approach wasn't stable.
The last plane that landed was slow to get off the runway, and the pilot or ATC decided that landing would create a risk of collision.
There was a plane or other vehicle on the runway that wasn't supposed to be there.
They encountered low-level wind shear.
The weather was bad, and they didn't have the runway in sight by the time they reached their minimums.
A passenger suddenly got airsick and had to get out of their seat to run to the restroom.
There are probably more that I can't think of right now.

Unfortunately, listening in to ATC broadcasts is illegal in England, so even if the tower did ask the reason for the go-around, it doesn't help us. There's also probably no publicly available record of the reason, so guessing is about the best we can do.

Personally experienced the passenger out of seat one, as soon as they said we were on approach and to remain seated, the guy gets up to go to the bathroom.  Even with the flight attendants yelling to him to stay seated, he just casually walked to the bathroom.  What's worse is that it was a very small airport, single aircraft stationed there, we were on the ground and out of the aircraft pretty quick, he could have waited 5 minutes.

@rtaft Some things just can't wait - vomiting is something that is very hard to repress.  On the other hand, a toilet break can generally be repressed for a bit.  Or maybe they're just selfish.

@Criggie. Could be deaf too, idk, he wasn't rushing and didn't turn around to look at or respond to the yelling flight attendant.  The timing of it, casualness, and lack of response made it seem intentional.

I've had to go to the bathroom so badly that I couldn't have waited five minutes. But I wasn't casually walking to the bathroom; I was running as fast as I could.

It is not at all unusual for commercial planes to do a "go-around" if the landing picture is not to professional standards.
Remember that a commercial plane frequently has between 100 and 250 lives on board.   There is simply no good reason to risk all those lives on a landing that is even a little bit out of "standard".  A go around costs a bit in terms of fuel, and maybe 20 extra minutes of time.  But compared to the safety of the passengers, it is not a concern.

There are plenty of YouTube videos of planes going around during gusty and windy conditions.
Sometimes birds or wildlife spotted in the area of the runway can be cause for a go-around, because hitting an animal during landing can be a serious problem.
Sometimes timing on the ground is an issue: if the prior aircraft doesn't quite clear the runway fast enough, or a ground vehicle is crossing the runway too slowly, its better to go around than to have a close call.
Sometimes the pilots' just have not established a well stabilized approach, due to workload, miscommunication, or difficult conditions.

It is much more rare (but not unheard of) for there to be a runway incursion, such as a vehicle (airplane or GSE) crossing the runway without permission or the pilots lining up for the incorrect runway.
I've spent a lot of time near the departure end of SeaTac airport, and can always hear when a plane suddenly spins up their engines for a go-around.  It probably happens to at least one flight every couple hours.  A single go-around is not a safety problem. (If a pilot has a bunch of repeated, unjustified go-arounds, they'll likely be sent for more training).  A go-around is far more preferable than the possible alternate outcome.


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Do helicopter rotor blades speed up and down during the rotation?
Does the inertia force change the angular velocity due to increase-decrease in the radius when the blade travels up? Does this cause an excessive bending moment?

This is what I understand that the question is, please feel free to roll back or edit otherwise if I did not get it right.

A helicopter rotor usually has flapping hinges, which lifts the blade up when travelling forward and down when travelling backwards. The blade being at an angle indeed causes a change in effective moment of inertia: it would want to speed up when at an incline, and this would indeed cause a bending or warping moment were it not for..

..the Drag Hinge or lead/lag hinge. The image is from this site which contains an explanation on other rotor heads as well.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


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Is it useful for twin propeller airplanes to have twin vertical stabilizers, in line with the propellers?
In some twin propeller airplanes, there are two vertical stabilizers which are on the same line as the propellers.
Is this shape helpful in terms of efficiency?
Like this one (twin boom tail):

I'm not really sure what you're asking. For example, the DHC-6 is a twin-engined prop plane with a single vertical stabilizer which is in the middle. The plane you've pictured is not a typical design so, if you're asking about twin-engine planes in general, it's not a good choice.

Welcome to Stack Exchange! I think I understand your question, and I've edited it a bit to try to make it a little clearer. If this is not what you meant, please feel free to change it back. Thank you!

A better question would be, could this configuration be helpful to efficiency? Maybe if the vertical surfaces were designed to extract some of the energy left in the swirl of the propeller slipstream...

Typically, vertical control surfaces were used to avoid have one large control surface. Twin tail planes like that pictured are pretty uncommon when compared to all the single fuselage/single tail planes with an engine on each wing.

The vertical tailplane is dimensioned as a tail volume: total vertical area times the moment arm to the CoG, to tackle the following issues:

Control after engine failure.
Lateral Stability.
Crosswind landings.
Spin recovery.

The vertical tail is subject to side forces, and the aircraft structure must be able to withstand the resulting reaction forces. The fuselage is very suitable for absorbing these by applying counter-torque, and is the go-to place for mounting the fin.
A large, long fin has a relatively long moment arm, and it can be beneficial to divide the vertical tail into multiple surfaces, provided there are suitable mounting spots. And behold, the P-38 had two tail booms and a horizontal connection bar, and the natural mounting method is two fins each at the end of a boom.
The fins are indeed in the prop wash and this does help their efficiency - in straight flight. At a sideslip, large angle of attack or in a spin, the prop wash will be directed away from the vertical tails, and these will be dimensioned as if there was no prop wash at all.

What about p-factor?

Yes it is, but many designs are more stable with one large V stab in the middle.
Some fine Navy engineers took the B-24 Liberator and turned it it the Privateer, making long flights over the ocean easier on the pilots due to improved stability.  The large V stab also plays an important role in avoiding "Dutch rolling tendencies", and, with its accompanying larger rudder, controlling side slip.
Putting control surfaces in the prop blast is an early form of directional control from thrust "fluid" flow, also seen in the V-2 rocket.  Smaller fins in stronger flow, it works.
So a designer could consider all three, as in the Constellation, remembering Rule 1:
FLY IT AS A GLIDER FIRST!


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What are the basic principles of making fighter jets manoeuvrable
Disregardimg thrust vectoring and other computer-aided functions, what are all the basic factors of making a fighter jet manoeuvrable? Using basic physics design principles such as high thrust to weight ratio, larger elevators etc., and without complicated programming and systems and changing of thrust directions?

This question is far too broad, there are entire books about maneuverability. Furthermore there's no way to explain the principles without complexity.

@GdD Sorry I beg to differ. Voting to keep open.

Info below is from TU Delft lecture ae4-211, Aerodynamics of the Fighter Airplane, Prof. ir. E Obert.
Fighter jets manoeuvre both in the vertical and in the horizontal plane. At the beginning of WWII dog-fight tactics were a leftover from WWI, when power loading was low and did not differ much between aircraft types, and diving speeds were limited and could be dangerous. Mainly defined by minimum turning radius and maximum roll rate.
In the later WWII years higher power loadings and -maximum diving speeds added exchange of potential and kinetic energy. The fighter generation developed in the 1970s (F-14, F-15, F-16, F-18 Mig 29, Su-27) had:

High thrust-to-weight ratio (T/W ≥ 1.0 in combat configuration)
Low wing loading
High maximum useable lift at high subsonic Mach numbers (wing-fuselage strakes, close-coupled canards).

Resulting in high load factors (up to 8g), high roll rates (up to 300 deg/sec), high angles of attack (30° steady state, >90° transient)


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Does $C_d$ change with speed?
In aircraft drag calculations, drag D is computed as $ D = C_D \cdot \frac{1}{2} \rho V^2 \cdot S$, with $C_D$ a dimensionless factor determined by the "streamline shape" of the measured object.
The question is: is $C_D$ always completely independent of airspeed? Are there any circumstances within aviation where measured $C_D$ values are variable with speed?
The reason for asking was sparked by car aerodynamics: Porsche claims that the rear spoiler

decreases drag at lower speeds, as the spoiler increases static pressure at the rear window, thereby "pushing" the car forwards more.

increases drag (and decreases lift) at higher speeds.

This can only  be possible if $C_d$ is different for different speeds, since cars do not change their angle of attack. Do fixed objects change $C_d$ with speeds? Are there any data results from the aviation world that confirm different $C_D$ at different speeds?
Or
If wind tunnel test meassured decrease in drag with fixed spoiler at 100km/h,will spoiler decrease drag at all speeds up to 300km/h as well?

No, Cd doesn't change with velocity. On a plane it would change only with angle of attack. For a car it wouldn't change at all for a given Mach number and Reynolds number.

Independent of your question - I had in mind that the spoiler of Porsches is adaptive. So at low speed they retract it, and at high speed - it gives more downforce (reduces lift on the rear of the car), and produces more stability. That would obviously change the shape of the car...

@Carl Berger, yes it is adaptive, at 205km/h spoiler increase AoA,but my question is if spoiler is fixed and if at 100km/h reduce drag does at all speeds up to 300km/h reduce drag as well?

@stevederekson555 Re changes with speed since car dimensions are constant, and that causes a change in c$_d$ as well.

@PeterKampf if Cd is changing with speed then spoiler can  possibly reduce drag at 100km/h but increase drag at 200km/h,(compare to car without spoiler)?

This depends on the location of the flow separation point which also changes with speed (more precisely, with Reynolds number) and can be influenced by adding a spoiler.

Please note this site is only for aviation-related questions as described in the [help] (Note the caveat "Aerodynamics (related to aircraft)" in the on-topic page). You may like to take a look at [physics.se] which might accept questions like this.

$C_D$ is a form factor variable, constructed to be as independent as possible of all the factors that contribute to aerodynamic forces. Two dimensionless flow parameters do influence outcomes of aerodynamic measurements:

The Mach number, which quantifies compressibility of air.
The Reynolds number, which accounts for viscosity/inertia effects in the boundary layer.

On our Aviation Stack, most airflow considerations are relevant to long slender bodies at flight speed (between take-off and cruise speed), meaning the range of Re is reasonably limited. For a car this may be a different matter.
image source
Re can be illustrated with the golf ball effect: the wake around the golf ball without dimples is what happens at lower Re, where viscosity effects dominate. At higher Re the boundary layer becomes turbulent and can follow the contours of the ball better, which creates less pressure drag.

Flight at lower Re takes place in gliders, and there are measurement data on glider airfoils at various Re. For instance in this book, Low Speed Airfoil Data by M. Selig, figure above is from page 68. It clearly shows how $C_D$ values change with Re at constant $C_L$. Note that differences at higher values are very small, at lower values very large.

This can only be possible if Cd is different for different speeds.

Yes indeed, at lower Re with laminar boundary layer we'll find a different $C_D$ than at high Re with mainly turbulent boundary layer.
Cars are short bodies moving entirely in ground effect, and usually the frontal area is used as the reference for $C_D$ as the form factor variable. Why? Because short streamline bodies dominant aerodynamic drag factor is the pressure drag, including suction pressure from the separated flow at the rear. And this flow separation takes place only at higher speeds, especially with the gently sloping rear of the Porsche - which reduces drag, but (unfortunately for a car) also produces local lift.

Pics above are from this very comprehensive master's thesis and show how the spoiler usually reduces drag: by reducing turbulence in the wake, thereby increasing the local air pressure. The thesis also mentions Porsche, and the way that OPs referenced spoiler works:

Companies such as Porsche, Bugatti or Mercedes have been using different technologies for spoilers and trying to maximizing the efficiency of it by eliminating the side effects in low speeds and increasing the advantages on high speeds. One of the most commonly used features is to have a hydraulic wing style spoiler at the rear end of vehicle that raises or lower at certain speeds to maintain down force on the backside of vehicle or to create air brake. This feature has been used mostly for safe driving. Spoiler deployment operation is usually automatic. The software operates the spoiler and fixes it in the certain height depends on the vehicle speed but the driver through a button in the cabin can also operate it. For instance, hydraulic spoiler that has been used in Bugatti Veyron comes up at high speeds to hold the car on the road better by creating down force. When the car reaches 220 km/h (140 mph), small hydraulic spoiler deploys from the rear bodywork and a wing extends about a foot. This configuration produces substantial down force, provides up to 330 pounds in front and 440 in the rear [16], which helps holding the car to the road in extreme speeds.

I'd assume the base line is a certain CD, if i increase the CD, wouldn't  it increases drag compared to the base line over the whole range? the shape function remains the same ..

@koyovis, my question is: if spoiler is fixed, does spoiler decrease drag at all speeds up to 300km/h?

Hard to say for this case, have you got a reference for the Porsche statements?

Please don't answer non aviation question on aviation. Instead direct users to more appropriate stack exchange sites if you would like to answer them there.

@Jamiec The question is relevant to aviation, as referenced in the answer.

@Koyovis the question does not ask anything about aviation whatsoever. We pilots can make anything about aviation ;)  See also: https://aviation.meta.stackexchange.com/questions/4151/should-we-allow-questions-which-are-likely-answerable-by-the-experts-here-but-a and https://aviation.meta.stackexchange.com/questions/431/should-we-answer-non-aviation-aerodynamics-questions

Do NOT change the question to fit your agenda!

Read. The. Answer.

For this answer I have to speculate since I do not have the wind tunnel data which would be needed for a more substantial answer.
Why should adding the spoiler reduce drag at lower speed? The most likely reason is a changed flow separation over the car's rear area. I would expect that flow separates already at some point along the rear window if no spoiler is present. Adding the spoiler reduces the slope of the car and will shift the separation point to the trailing edge of the spoiler. This should indeed raise the pressure over the rear window, thus pushing the car forward.
Why does the spoiler add drag at high speed? The higher Reynolds number at higher speed should delay separation and cause attached flow over the rear window plus a suction peak at the higher curvature below the rear window before the flow separates. Adding the spoiler adds more wetted surface and a downward-directed lift force which will cause its own induced drag. But now the suction peak is avoided and some downforce is added instead. The separation occurs again at the trailing edge of the spoiler, so the separated area is a bit larger than without spoiler.
This small drag reduction is not the reason for the spoiler, though. It is really needed to avoid that suction peak at high speed which will make the car impossible to control, especially in crosswinds. Audi learned this the hard way when the designers first would not accept a spoiler on the Audi TT since that would had destroyed the symmetry of the car's contours. After a couple of crashes the engineers eventually could overrule them.

So in short, re "Do fixed object change Cd with speeds?", the answer is "Yes, due to changes in flow separation point, which are entirely due to the changed Reynolds number"?

@quietflyer So in short: Yes.

@PeterKampf can post link of Audi problem

@LostincurvedSpace-Time Sorry, I omitted which car that was. Here would be one article about it. It even reduced drag at high speed a bit.

Please don't answer non aviation question on aviation. Instead direct users to more appropriate stack exchange sites if you would like to answer them there.

The drag coefficient is designed to be a coefficient - and hence be as independent from the flow state as possible.
Yet, the simple cd definition $c_d = D / (1/2 \rho u^2 \times S) $ just uses the (incompressible) dynamic head of the incoming airflow as reference pressure. This way effects of viscosity are neglected, as are effects of compressibility. As example, there are the compressibility effects over Mach number, a classic chart would look like this:

To be fair -  neglecting compressibility effects up to Mach 0.3 leads to errors below 5%, so for a Porsche at 300 km/h (which isn't far from Mach 0.3), that should be fine.

Yes, if only because Cd changes with the Reynolds number, and –all other things equal– the Reynolds number is a function of the airspeed.

Effect of Reynolds/Mach changes at usual speeds of a car is negligible. Even for planes it usually isn't considered in simple analyses.

+1: it's true, in normal flight range $C_D$ change with Re is negligible, with cars & gliders it is a different matter though. The answer is a bit brief though.


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Is fuel efficiency the only reason why hypersonic missiles don't use rocket engines?
Is fuel efficiency the only reason why hypersonic missiles don't use rocket engines?
And why are rocket engine missiles relatively slow compared to scramjet missiles? In theory rocket engines can produce more than enough thrust for 24Mach speeds?

Any Department of Defense requirement that needs a hypersonic missile to fill a strategic or tactical mission role will not have fuel efficiency at the top of the list of design criteria.

@MichaelHall but fuel efficency leads to short range..

Liquid propellant rockets are quite complex to keep in readiness, much more so than scramjets.  If you’re going to the trouble of building an ICBM then you’ll factor this in, but for a relatively mass-produced device an air-breathing engine is much more attractive.

I think most of the hyper-sonic missiles do use rocket engines. Remember a hyper-sonic missile, in practice, is a solid-motor ballistic missile with a waverider warhead. To accelerate to hyper-sonic speed and climb to that altitude you have to use a rocket booster, but once you are there however, the a cruise engine isn't entirely necessary.

Not an answer but useful data point https://en.wikipedia.org/wiki/Sprint_(missile) which was a mach 10 pure rocket with a starting mass around three and a half tonnes for final payload around 500kg. flight time of 15 seconds to around 40km. Not a particularly useful platform for a cruise missile.

Actually inefficiency would lead to a shorter range, but I get your point.  I'm looking at this from the perspective that efficiency = economy, often resulting in a compromise of max power.  My point is that you design a missile to meet a certain mission profile and while the type of fuel it consumes certainly play a part in the max range, fuel economy, (admittedly my paradigm based on the terminology...) isn't as important as getting the job done.

Hypersonic missiles fly through the atmosphere and remain manoeuvrable until impact, making it much harder to defend from them. And since they're flying through the atmosphere they use atmospheric air for combustion.
Pic source
The main advantages of rocket engines:

They don’t require air to maintain combustion, so they can fly trajectories that take them outside of the atmosphere while still providing thrust.
Since they carry everything they need on board, they can be launched from practically anywhere, even from submarines.

The main disadvantage of rocket engines:

They need to carry their own oxidiser, thereby reducing the maximum fuel weight. The Tsiolkovsky equation teaches us the extent of limitations of payload as a proportion of empty weight and fuel weight - but a large enough rocket can still carry men to the moon.
Long range rockets need to employ multiple stages, ejecting bits of empty fuselage weight with its engines, and continue flight with other engines.

Pic source
Hypersonic missiles only need to carry fuel and payload, which they oxidise with air which is also used as a reaction mass. Without having to eject bits of its construction with perfectly functioning engines, and using wings to create lift and flight control surfaces to control flight path just like an aeroplane. But launching them is an issue: flying faster than Mach 5 they are propelled by scramjets, which cannot be operated from a standstill like rockets can. So they need to be launched by either a rocket or from an aeroplane, preferably a really fast one.

Is fuel efficiency the only reason why hypersonic missile don't use rocket engines?

It's the main reason, yes. While exclusively flying trough the atmosphere it makes no sense to not use the unlimited amount of reaction mass of atmospheric air for combustion.

And why are rocket engine missiles relatively slow compared to scramjet missiles? In theory rocket engines can produce more than enough thrust for 24Mach speeds?

Whether rocket propelled or scramjet propelled, when flying through the atmosphere the missile heats up and therefore there are limits on the maximum attainable velocity. A rocket propelled missile can definitely reach the same speed as a hypersonic missile, there is nothing slow about them.
The two pictures linked to a from a site that explains the basics of missile threats, a worthwhile read.

You mentioned reaction mass shortly, but maybe you could repeat that in your summary paragraph starting with "It's the main reason, yes." Basically, the value proposition of a jet engine over a rocket motor is "Free, unlimited oxidizer and free, unlimited reaction mass". Even ion engines which don't need fuel or oxidizer are still limited by their reaction mass.

Mach 24 in the atmosphere? Not for long.  Ask any meteor.
Developing airbreathing technology may improve fuel efficiency, but oops, the drag of the intake and heating from compression may cancel any advantages airbreathing technology gives.  In short, try as they may, hypersonic jets do not seem to be advancing beyond the research stage.
The Kh-47M2 Kinzhal air launched solid fueled rocket is a very practical solution, but one that could be answered with one of these, also carried by an aircraft, if needed.
Speed vs maneuverability rages on.

It is my sincere hope that aviation research continues for $peaceful$ purposes.

@Roberto So why they claim that missile speed is around 20mach?

@JurgenM to sell it.  Orbital and suborbital craft commonly make that speed outside the drag of the atmosphere.


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Why force the nose of 737 Max down in the first place?
Can someone here explain a pretty basic question that I have had  ever since this controversy came to light? I don't understand why any aircraft manufacturer, engineer, software developer, would make a device that forces the nose of the plane down. You can't land unless you're in control and landing safely is the only reason to go down that I know of. So I just don't get that as an initial thing, but the second part of that is, given the hazards of pointing the nose down, which are obvious even to me, why would you not design the system so that it's a warning, and allow the pilots who are actually sitting there to decide whether going down is a good idea or not? That just fundamentally, as a threshold issue, makes no sense to me at all. So I'm hoping you could help me understand. Thank you.

I've heard it's so that the MAX can have a common type rating with the older 737s, which airlines like in order to keep training costs down

The main thing to avoid in aeroplane stability & control, is an aerodynamic nose up moment that is not commanded by the pilot. The uncommanded nose-up moment would not auto-stabilise, but rapidly get progressively larger with increasing angle of attack, and run away to a stalled aeroplane.
During certification of a passenger aeroplane, many tests are carried out to check if the airframe does not start to have a mind of its own.

If the pilot does not provide a control input, the airframe must return to the trimmed position.
Forces and inputs to move the aeroplane away from trimmed position must be such that there is an ever increasing force required to achieve an ever increasing nose-up position. The nose-up position must always be commanded by the flight control surfaces, elevator and stabiliser, in a predictable way.

One of the tests to be performed during certification is stick-force-per-g. Bank the aeroplane and start turning while pulling the stick back in order to maintain altitude. Then bank more and pull back more, in ever tightening turns. It must be progressively harder to pull back on the stick to maintain altitude, never easier.
It was during this wind-up turn that due to the engine configuration of the MAX an aerodynamic nose-up moment appeared, which would cause the pitch stick force to suddenly become less than expected. Not as bad as a runaway pitch, but still an undesirable situation when the pilot is still straining to maintain the manoeuvre. This is the situation that MCAS was originally designed for, to auto-compensate for this situation only. The aeroplane should always stabilise itself and manner of control must be predictable and within the reaction times of humans.
More info in the links in this question, very interesting articles which illustrate how MCAS design bloated from the original scope, under time pressure.

The accelerated stall.  They keep blaming the engine (could be part of it).  Yet using smaller "more efficient" airfoils as stabiliators has one very troublesome aspect. With a flat plate, the coefficient of "lift" is linear to the angle of deflection all the way up to 45 degrees, along with loads of drag, making the pull harder and harder.  Notice the lift curve on an airfoil is anything but linear post stall.  Could this be part of the control issue?

@RobertDiGiovanni Not a stall, but a tight turn. I could not find any reference to a change in stabilator profile relative to the 737NG.

Well, I am going to award the check mark to Koyovis, because his answer is the most detailed, albeit a little technical for me. But big thank yous also to Anthony X and Alexei for less technical answers that rounded out my understanding. Even the comments were helpful!

The second sentence of the first paragraph appears to describe a situation in which an airplane becomes statically unstable in pitch (...would not auto-stabilize...run away to a stalled aeroplane.) As that issue has been  raised in this answer, I think it would help to point out that the problem here was not that bad -  the airplane does not become unstable (at least prior to the stall), but its handling characteristics, unmodified, were still unacceptable.

What about the idea that changes to the MAX reduced g-break (nose drop) at stall, so MCAS was added just to simulate g-break via pitch trim, and fatally could apply it repeatedly as the pilots went in and out of MCAS mode while troubleshooting?

@ami It sounds like an assembly of half-truths to me.

"The main thing to avoid in aeroplane stability & control, is an aerodynamic nose up moment that is not commanded by the pilot.".. funny. I would have though the main thing to avoid in aeroplane stability and control is flying underground.

Counter-intuitively, lowering the nose of an aircraft is not done for the purpose of "going down". Climb/descent is managed with throttle, and speed is managed with the control column/stick. The logic of this makes sense when you consider that going up or down involves the addition or removal of potential energy which is sourced by the engines and sunk into drag.
Pull back on the column/stick and you increase pitch angle, which increases angle of attack, which increases lift. That causes the aircraft to go up, yes, but going up requires extra energy which most immediately comes from the aircraft's forward motion - it slows down. As it slows, lift decreases and (ideally), equilibrium is restored at some lower airspeed and increased angle of attack.
Same thing in reverse - push forward and you go faster.
Due to the way the new engines were fitted to the 737 Max 8, it acquired a handling problem in which the nose could pitch up unexpectedly; if uncorrected, this pitch-up could induce a stall. MCAS was intended to compensate for this bit of bad behavior by detecting its occurrence and automatically pushing the nose down to maintain the expected attitude and prevent a stall. The crashes tragically revealed points of failure in the MCAS; in those cases, MCAS responded to a falsely indicated high angle of attack condition.
A fairly lengthy article was published at The Verge recently, which details the story of the 737's evolution into the Max 8. The article outlines the reason the MCAS exists and why it does what it does (both nominally and erroneously), as well as a short history of one of the accident aircraft leading up to its final moments. Not sure how authoritative the article is, but I believe all the stated facts are correct.

To add a bit to the existing answers, the reason for the unexpected pitch-up moment on the 737 MAX, as far as I understand it, had to do with the flattened portion on the bottom of the engine cowling.
The root of the problem is that the 737 was designed back in the day of low-bypass turbofans (specifically, the Pratt & Whitney JT8D.) Due to the low bypass ratio around the core, these engines had much smaller diameters than today's high-bypass turbofans. The JT8D had a fan diameter of only 49 inches, while the LEAP-1B on the 737 MAX has a fan diameter of 69 inches (and even that is notably held smaller than the 78 inch diameter on the LEAP-1As for the A320neo series.)
Due to the massive increase of engine diameter with the switch to high-bypass turbofan engines, ground clearance became a problem, since the landing gear height was designed with much smaller engines in mind. So, you will see even on the 737 NG series that there is a flattened portion on the bottom of the engine cowling in order to improve the ground clearance a bit while still allowing the landing gear to fit into the wheel wells. Since the LEAP engines on the MAX are even larger than the CFM56 engines on the NG, the flattened part became larger in order to fit the new engines under the wing.
737NG with flattened engine cowlings (Source)
It turns out that this flattened portion can create a significant amount of lift (and, due to its positioning, nose-up moment) at high angles of attack. This causes the aircraft to want to pitch up even more at high AoA, which is bad for the reasons that Koyovis has already explained well. Thus, MCAS was designed to prevent this tendency for AoA to continue uncommanded increases at high AoA by intentionally pushing the nose down if the AoA got too high.
The idea itself is not necessarily a bad one and systems with similar purposes (pushing the nose down to prevent AoA from getting too high) exist on nearly all Airbus planes still flying and also on other newer Boeing designs. The problem was with the implementation of the MCAS, which apparently did not cross-check the AoA vane inputs against each other or otherwise do sufficient sanity-checking on the inputs from from the AoA vanes before acting on one of the AoA vane inputs to push the nose down without command from the pilots. In the case of the Ethiopian crash, preliminary information from the investigators indicates that the AoA vane that the MCAS was using for input was likely completely sheared off of the aircraft, possibly during a bird strike or similar FOD event during takeoff. Due to the counterweight of the vane still being attached, this caused it to indicate an extremely high AoA, which in turn triggered the MCAS.

Source for the profile of the engines creating lift, please? The origin of the pitching moment is the moving of the thrust-line of the engines in the MAX 8 away from the centre of gravity. This was necessitated by the use of new larger, more efficient, high-bypass turbofan engines. In a throttle-up scenario (eg taking-off) the aircraft now pitches up more than its predecessor version. To reduce pilot training costs for the new type, MCAS attempted to silently reduce the pitch moment perceived by the pilot, by using the horizontal stabilizer to apply a corresponding pitch-down moment.

...this, together with a number of other design and communication missteps (trim wheel size reduction, failure to fully reset trim MCAS parameters in a restart, lack of redundancy in AOA sensors, failure to mention the system in the iPad-based type conversion training) led to a a hole through the block of cheese, delivering two disasters.

@Ben High power settings do increase the pitch-up moment further, but that's not what MCAS was designed to counter and MCAS isn't active during takeoff anyway (it's locked out when the flaps are selected.) See here for a more detailed description. It's the nacelle shape that causes the lift generation. If it were just the thrust, it would apply to virtually all airliners these days, as the engines are almost always mounted well forward of the CoG and create pitch-up moment at higher power settings.

Thanks. So strangely it is the repositioning of the engines and the thrust they generate but not from the turbofan, but from the fairing dynamics?

@Ben It's both the nacelle shape and the position of it (and the combination of the two.) Normally-shaped nacelles wouldn't generate as much lift at high AoA as far as I know. Of course, the farther forward you move the engine, the larger the pitching moment will be for a given lift. In this case, both the lift increased due to larger and  reshaped nacelle and it was moved a bit farther forward.

Note that the flattened engine cowling profile was introduced in earlier variants.

@Ben Indeed, as noted in the answer. It was expanded from the NG to the MAX, though, in order to compensate for the LEAP engines, which are larger than the CFM56s on the NG. The lift at high AoA existed in the NG, but it's larger and a bit farther forward in the MAX, resulting in more pitch-up moment than the NG nacelles caused.

Thank you. To clarify: are you saying that the moving of the turbofan thrust-line is not causative in this instance? You are saying that it is solely down to the passive aerodynamics of the nacelle?

@Ben Yes, pitch up moment from thrust was not the reason MCAS was created.

As a passenger who doesn't build aeroplanes, or fly them, but who does program computers - I am unnerved by the idea of building something that basically doesn't fly well, and then just getting computers to step in and sort out any unexpected lurches...

@GrimmTheOpiner As a pilot and embedded systems software engineer who also spent some time working at an aerodynamics testing facility, it generally flies fine, but had one relatively small edge of the flight envelope (where you really shouldn't be in normal operation anyway) where it didn't have the proper longitudinal stability. So, putting in a software fix for that isn't necessarily a bad idea. It was the design and implementation of the MCAS (and its lack of communication to pilots) that was just awful and completely baffling as both an engineer and a pilot.

MCAS forces the nose of the 737 MAX down because, under some circumstances, the nose of the aircraft can pitch up and this may result in a stall.
From Wikipedia:

The Maneuvering Characteristics Augmentation System (MCAS) is a flight control software system developed for the Boeing 737 MAX to provide handling qualities similar to the Boeing 737 NG, especially in low-speed and high angle of attack (AoA) flight. It lowers the nose without pilot action when it determines the aircraft is too nose-high, based on input from airspeed, altitude and angle of attack sensors. However, it is susceptible to erroneous activation, as evidenced in the deadly crashes of Lion Air Flight 610 and Ethiopian Airlines Flight 302. The 737 MAX is indefinitely grounded until regulators decide the aircraft is airworthy, pending software and instrumentation updates and revisions to information for flight crews. They may also be required to undergo MCAS training sessions in flight simulators.

Reasonable answer, but yet again you are using wording that appears to have led to the original question: MCAS is NOT supposed to push the nose down, it IS supposed to push the nose lower than it is - the assumption being that the nose is too high; MCAS should make it slightly less high.

@MikeBrockington: What's the difference between lower and down? A 737 isn't going to be flying inverted, so the two appear functionally equivalent.

@MSalters - Both words refer to the direction of change, but 'down' can be misinterpreted as meaning 'below horizontal'

@RobinBennett: So can 'lower'. Perhaps the better description would be that MCAS is supposed to level off.

MCAS is not supposed to 'level off' either (as I understand the term). Whether the plane is approaching a stall while climbing or while diving (slightly), MCAS is merely supposed to reduce the climb-angle enough to avoid a stall, not necessarily to make the plane dive.

@MSalters  To a layman, 'Down' means below the horizontal, i.e. diving.  'Lower' means less far up, less of a climb; this may possibly mean a transition from climb to level, from level to dive, but not necessarily.

IIRC, the modifications to the 737 (including the new engines that powered it) meant the engines had to be positioned further forward and higher. (This was a similar case when they swapped from the original "cigars" to the CFM-56). And under some flight conditions, this resulted in automatic pitching up of the aircraft.
MCAS was designed to counter this tendency, to avoid pilots constantly having to make corrections. Since it was inherent behaviour of the new design, it was implemented in such a way that the pilots were not intended to even know such a system existed on their aircraft, for all intents and purposes it was flying "just as the other ones did".

I've seen this explanation elsewhere, but I would have thought that raising the thrust line of the engines would cause a nose-down moment, not nose-up? Maybe it was just that these engines were more powerful than the old ones?

@RobinBennett that's an interesting topic, maybe you should ask this as a new question.

The new engines are so big that they couldn't raise them far enough to maintain the original thrust line. They were raised such that part of the intake nacelle was already above the wing!

The issue with the 737 MAX is not pitch caused by the thrust of the engines (though that does exist, on the 737 and other airliners). The issue is that the engines with their large fairings generate a decent amount of lift at higher angles of attack; combined with the forward position of the engines that lift generates the nose-up tendency that MCAS was meant to counter.

For stable flight all forces and moments must be at equilibrium. For a given speed and thrust there is a window of acceptable angles of attack (AoA) when the wing produces enough lift. If the angle of attack is too high, the wing stalls and plane loses lift abruptly. The critical AoA can be reached by either excessive pitch up for a given speed or slow-down for a given pitch.
The Boeing 737 design dates back to 1964 and the latest 737s are evolutions of that design. The most apparent differences in design of the now-grounded planes are:

Hull lenght
Wingspan
Engine thrust

If you compare engine positions and the fact that the thrust doubled over time you can clearly see where the unintended pitch up is coming from.
This pitch up can lead to stalling and pitch down countermove must be performed to avoid such situation because power-up will lead to further pitch up move.
Also note that the MCAS intervention led to a disater only twice; in all other cases not worth to be documented it worked as intended. The issue, that made all 737-MAX grounded, is NOT about that such system was implementd but how it was implemented and how it was documented - which is a comletely different story. Also note that it is not the first case of anti-stall not-well-documented safety system malfunction leading to disastrous consequences.

Yes, not the first case indeed. Airbus has also had its fair share of problems with its alpha protection systems, too, including one incident that nearly caused a mid-air collision between an A330 and an A340 over the Atlantic from a dramatic uncommanded climb and another in which an A330 made two extreme uncommanded dives (one of which actually went into negative g territory) that injured several passengers and crew members before the pilots were able to get the airplane back under control.

“...led to disaster only twice.” Once too many.

@Koyovis Twice too many, I'd argue. Sound software engineering and a culture of airlines trying to make sure their pilots learn as much as possible about the immensely complicated machines they're flying rather than as little as they can get away with would have prevented both disasters.

I don't understand why any aircraft manufacturer, engineer, software developer, would make a device that forces the nose of the plane down.

Because history has proven that pilots themselves don't always do it.
I recon the engineers thought it would improve safety.
Looking at the Air France flight 447, if (a working) MCAS was in place it probably would have saved the plane.
In short the first officer stalled the plane in to the water.

First officer Robert said to himself, "climb" four times. Bonin heard this and replied, "But I've been at maximum nose-up for a while!" Captain Dubois realized Bonin was causing the stall, causing him to shout, "No no no, don't climb!"

This is just one accident where lowering the nose would have saved the plane.
I'm not blaming the air crew, it's "counter-intuitive" to lower the nose and once your reptile brain is in charge it is hard to get back in control of your mind.

Piloting 101: when in a stall, lower the nose. The pilot may do that.

@koyovis yes. They should do that but I believe AF447 is a clear example that it's not always done. Many other accidents has been close to ground and but AF447 had lots of time to solve the problem yet they didn't. Is a MCAS or similar needed, perhaps not needed but it could help.

It's worth consideration I guess, T-tail aircraft have stick pushers as well.

Planes fly with speed. If they don't have enough speed, they fall down like bricks.
Nose down trades altitude for speed. You're losing altitude, but you're "flying more".
Nose up tries to trade speed for altitude. At best, you're gaining altitude, but you're "flying less". At worst, you're not gaining anything, you're just slipping down while looking up. This is how Air France Flight 447 crashed: copilot kept pulling up and they've lost all altitude just by that.
737 MAX, due to it's unplanned engine size, has natural tendency of pulling the nose up, which is super-dangerous because of the reasons above. Pushing it down is a way to keep this danger at bay. The problem is not in pushing the nose down itself, but pushing it down too often and too much. They've replaced big danger with as smaller one - which unfortunately proven to be designed even worse than oversized engines.

The real problem is not that the lift from the nacelles provide a upward pitching moment.  The issue is why.  Stability demands that the center of lift of the total airplane be aft of the center of gravity.  Total lift comes from several sources, the wing, some contribution from the fuselage, contribution due to the engine thrust angle relative to the angle of attack, and engine nacelle contribution.  At high angle of attack the engine nacelle contribution becomes significant.  Add all of these sources and the result is that center of lift of the total airplane moves forward at high angle of attack.  As angle of attack is increased the center of lift moves forward resulting in a pitch up.  As angle of attack is progressively increased the center of lift moves forward of the center of gravity and potentially uncontrolled pitch up occurs.

I read an article about this that helped me understand it very well.
The engines built for the 737 Max are very powerful.
Think of yourself riding a motorcycle or car. What happens to the car nose level when you start at a fast rate? The front moves up. And when slowing down fast? The front moves down. Right?
Due to this problem according to the engineers, something needed to control the level of the plane which is what the MCAS was designed for. To help stabilize the plane during these quick movements events speeding up or quickly slowing down.
With correct data fed to MCAS, it works very well stabilizing the plane throughout the flight so the pilot can fly the plane like a 737 as it was designed to be.
Without MCAS, the pilot would have a difficult time keeping the plane stabilized when the speed of the plane shifts. Pilot can't focus on many things going on without MCAS.
That in a nutshell made alot of sense to me. Any plane with powerful engines is going to need a MCAS type of system for this reason. And it was designed to run in the background.
Driving our cars, lots of devices are at work in the background enabling us to drive and focus on the road. Same principle here for the pilot because of the new powerful engines.

I think the issue for Boeing was the risk of falling behind Airbus in sales. So they modifed the 737 using far larger engines which because of limited ground clearance had to be mounted in such a way that increased the risk of stall during take off. So with a plane that is not as aerodynamic as it should be Boeing fitted a system to overcome the aerodynamic flaws, MCAS. It seems strange to knowingly produce a plane which is not as aerodynamic as its predecessor, but in their desperation to keep up with Airbus, Boeing did exactly that.

Also the regulator allowed it happen

There was indeed pressure on Boeing to come up with a competitor for the Neo, and the limited ground clearance made things difficult. It's important to state that there seemed to be nothing wrong with the original design of MCAS: automatic correction in a very specific circumstance that pilots would almost never find themselves in. The design scope had to be expanded in a hurry though, and that is where the problems seem to stem from.

This had nothing to do with takeoff. The MCAS is not even active during takeoff. It's locked out when flaps are selected.


text:
Why does my Cherokee PA28-180 Lycoming motor run hot?
Regarding my Cherokee PA28-180. Could you please explain me why my new Lycoming engine with 800 hrs runs hot?

My mechanic says that the EGT #2 should be around 1350 degrees. The minute you lean the fuel at cruise alt, you have to richen it back up to stay < 1350 degrees.
During 1800 rpm run-up & during cruise speed of 135 mph, the oil pressure is just barely under the red line & it stays there for the duration of the flight. My mechanic says that's ok for this motor.
My mechanic says that we are to stay away from the 2100 rpm to 2400 rpm range. That's red line for our motor. Why?

also I've been leaning using the lean-rpm drop-richen back up until motor is running smooth. How do you lean utilizing the EGT gauge? My mechanic says to fly w/ the #2 cylinder <1350 degrees.

I'm voting to close this question as off-topic because we aren't mechanics and, even if we were, we couldn't give advice on an engine we've not seen.

This guy has a question that can be answered.  The answer isn't a matter of opinion and can be applied to virtually every common aircraft engine.  How is this too broad??

You say it runs hot.  Why do you say that?  Is the oil temp pushing redline?  Do you have a CHT gauge?

EGT alone means nothing. Don't use it to set cruise settings.  What if the probes aren't all set 5" (or whatever) from the exhaust valve? What if they're only 3"?  Or 7"?  Or staggered between 3" and 7"?  Point is, the EGT number is just a repeatable value.  Don't think of it as a temperature.  Your Archer with probes 5" from the exhaust valve may read 1350° at full rich takeoff from sea level.  Your hangar mate with an identical plane (except for having probes only 3" from the valves) might see 1420° at the exact same setting.  Once you understand that EGTs aren't really temperatures and are merely reference numbers (or some type of "heat unit that isn't 'temperature'") the whole thing makes more sense.
Also, if you just have a single point EGT then just pretend it's not there and simply use the POH.  Single point EGTs tend to do more harm than good, especially if they're reading the confluence of two cylinder exhausts.  What possible useful information can you glean from an averaged temperature from half the cylinders on your engine, especially if it's carbureted?!?!

Does your mechanic know (and can he explain) why his tips make sense?  I'd love to hear his reasoning because I don't think it'll make sense.

You either have a bad oil pressure gauge or your pressure it too low.  Fix that.  A good gauge running below red line is very bad.  A bad gauge is better, but still bad.  The limits are not arbitrary.

Engine RPM limits are important.  Does Lyc publish a limitation between those RPMs for your specific engine suffix and serial number range?  If the limitation is an AD, has it been complied with and the limitation removed?
Running your engine at redline will not hurt anything provided the engine does not have a limitation (in the limitation section of the POH) on max RPM.  I know it feels right to "baby" the engine but there is no practical effect unless a limitation exists.  Note that "recommended procedures" are not "limitations."

I'd seriously consider talking with another mechanic if that's at all practical.  Your mechanic may be an airframe god but his advice regarding engines would have me shopping for a new IA.

so you have two very good questions:
1)1350 egt, look up the egt limits from Lycoming for the correct engine model.  What you should be looking with most non-turbo charged engines from an operations stand point would be - set cruise power, very slowly lean to peak and then enrichen the mixture until you are 50-100 degrees rich of peak. Please be aware that leaning the engine is something that is condition specific (OAT, Density Alt, Power Setting, IAS) Changes to any of these can and will change your EGT.  This will typically give you a safe margin and not have the engine detonating or running too hot.  I have not seen a Lycoming where the egt at 1350 is outside of normal operating temp however in some conditions 1350 could be pushing you close to peak.   Check with Lycoming on your model engine just to be sure.
2) RPM is not limited to 2400 on the o-360 in the Piper PA-28.   RPM range is limited on some of the year models where you are not supposed to run the airplane in the 2100-2400 range due to a vibration issue with the standard propeller / engine combination. The performance section of your POH will give you % hp verses RPM at specific altitudes.

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text:
Can a supersonic aircraft have winglets?
In order to have laminar flow over the wings which helps in friction drag reduction, a number of airfoil designs have been proposed from the past to the latest 737 winglets.
But if we design a laminar flow wing for supersonic aircraft, the wing should be very thin which will end up with stall and speed related problems.
So, I am curious that if we add winglets on supersonic aircraft, will it be helpful other than for its drag? If so what are all the problems we need to consider?

What initial research have you done, casually or formally, before asking this question?

XB70 has a pretty big "winglet", and Mig-31D(anti-satellite variant) has a small one. But those are more for vertical stability (effectively increasing the area of vertical stabilizer).

@KorvinStarmast. Actually when I was doing some research on cranked delta wings for the Supersonic Business Jet, I came across the word "Boeing 737 winglets". Thats how I ended with this question.

Could you please summarize that and put it into your question?

This has been (and thus can be) done and there is some current research on it as well. The XB-70 Valkyrie had dropping wing tips that added stability at high Mach speeds as well as allowed the plane to ride on its own shock wave. It also effected trim drag,

The repositioned wingtips also reduced the area behind the airplane's
  center of gravity, which reduced trim drag. The downturned outer
  panels also provided more vertical surface to improve directional
  stability at high Mach numbers.

You can check out this podcast episode that interviews some of the people at NASA doing various research projects including wingtip control systems and other interesting applications in that space.
You may find some info in this book on supersonic wing theory.
This thread (and you know how accurate the internet is) has some interesting points on the matter.

Research has been carried out into this area, for instance the Computed Fluid Dynamics tests that NASA carried out on highly swept wings in supersonic flow. They reached the conclusion that winglets can have positive effect on lift/drag, but it is of course highly dependent on the wing shape and profile. From the report:

winglets can be designed and aligned in supersonic flow such that little or no performance penalty will be incurred relative to a wing of equal projected span.
alteration of winglet length, sweep, and camber may not be as critical as the toe angle or the orientation of the winglet relative to the wing.
winglets with negative dihedral and toe in were found to typically outperform similar winglets with positive dihedral.

One of the configurations studied is below: NACA 1402 base wing with 65 deg leading edge, and the best winglet found. One of the recommendations made is to put further research into dual use of winglets, if they are useful for yaw control at supersonic speeds.

The projected European spaceplane Hermes did indeed have winglets. Those were needed for drag reduction in the landing phase. So they did not help at supersonic but at subsonic speed.
Without the winglets the energy would had bled off too quickly during rotation so the descent could not be stopped in simulations. The spaceplane without winglets was in essence impossible to land. Only by lifting the subsonic L/D above 5 with the help of the outward canted winglets did the simulation result in smooth landings.

Artist impression of Hermes spaceplane (picture source)

Even if they can, would they be practical? Many combat aircraft use the wing tips to carry short range missiles or other light weight stores.
You'd lose those hardpoints.
Plus you'd introduce more complexity into the wing, especially if you rely on them for stability, making the aircraft more vulnerable.
Also, I've not done the calculations but many combat aircraft have wings that have a large taper, winglets may very well be only marginally effective with such a wing design.

But how about in case of Supersonic Business Jet other than fighter jets?

@SathiyaNathan the effectiveness of winglets depends at least in part on the wing design. A long narrow wing is what we see winglets used on most often. If your business jet had such, they might work, especially during the subsonic phase of its flight, which would probably be a lot of the flight profile as most countries prohibit supersonic overflight except over open bodies of water (sea, maybe very large lakes).

There are 3 types of drag:

Friction drag
Induced drag
Wave drag

In low speed flights, first two are responsible for the total drag created. Wave drag is usually neglected as it is very less compared to others. Winglets are provided to reduce the induced drag.
Now, in case of the supersonic aircrafts, wave drag creates most of the total drag as the shock wave creates enormous drag. Friction drag and wave drag are proportional with the velocity, and induced drag is inversely proportional to the velocity of flight. In case of supersonic flight, due to the velocity, the induced drag is going to be very small that it is always neglected.
So yes winglets are possible on a supersonic aircraft, but are rather decoration then useful. And most supersonic aircraft are military fighter jets, and fuel efficiency in the military is not a priority
By the way, this answer is from a Quora question found with a quick Google search.

Are induced drag and wave drag not normal?

Let me change it to General dag

Is he a high ranking military officer? A specific term would be more appropiate, like friction drag.

I will change it, thnx for the tip ;)

But anti-satellite variant of Mig-31 (and XB70?) does have winglets. So what about other use cases than drag reduction?

I bet those are for meant for other things (Stealth?)

"There are 3 types of drag" - Four, actually (you forgot pressure drag).

@HugoWoesthuis They do increase lift, by keeping pressure waves confined under the wing.

optimizing the wing for wave drag (=shorter wings) increases induced drag. By needing larger wings, it then increases drag again. If we remove the induced drag, we have more freedom to use a more wave-drag optimized wing, reducing drag even more.

The North American XB-70 Valkyrie was a Mach 3 bomber which had variable-geometry winglets. Their purpose was most unusual in that they turned vertically down to help reduce drag during supersonic cruise, and returned to horizontal to increase the wing area for takeoff and landing.
The wing of the Valkyrie was a waveriding delta. The engine intakes formed a central underwing body which created a strong shock wave. The wing leading edge was angled such that at cruising speed the shock front ran out along the wing just behind it. All the wing behind it was subject to an increase in air pressure, creating a substantial amount of extra lift and improving the lift/drag ratio of the wing. When the shock met the downturned winglets, it was reflected back under the rear of the wing creating even more lift and improving cruise efficiency even more.
The winglets also provided additional directional stability just when it was needed, as this normally reduces at high speeds.